This invention relates generally to an advanced integrated propulsion system (AIPS) for gas turbines, and more particularly, to structural improvements that improve the efficiency of the gas turbines for propulsion or energy.
In general, a propulsion system must have the highest efficiency at variable load, and especially at partial load, in which a vehicle is used for the majority of its life time.
The very high and dramatic variation of the specific fuel consumption is, in fact, the major disadvantage of the gas turbine engine and the Brayton cycle it follows. The gas turbine is, therefore, generally unsuited for applications where power and speed are variable.
Therefore, the gas turbine has ordinarily been used only in applications where the requirements for power and speed are substantially constant.
The superior characteristics of the gas turbine engine for numerous high power applications, both military and commercial, are well known. Low specific weight and fundamental mechanical simplicity particularly favor this type of engine.
Numerous developmental paths are available for reducing the gas turbine engine's fuel and air consumption and for reducing its size and weight. Many of these paths, however, lead to undesired complexity and high cost.
One obstacle in the evolution of the gas turbine engine is the inability to provide a sufficiently high and constant pressure ratio for the compressor without unduly increasing the engine's size and cost. The effects of the compressor's pressure ratio and efficiency on the engine's specific fuel and air consumption are dramatic. Although an increase in the pressure ratio at a given efficiency increases the engine's specific power, a more significant effect is the reduction in specific fuel consumption.
Another obstacle in the evolution of gas turbine engines is the limited heat resistance of the engine's fixed vanes and turbine blades. Complete combustion of petroleum fuels at a stoichiometric ratio with air results in a combustion gas temperature near 4000.degree. F. or 2200.degree. C.). However, the engine components mentioned above cannot ordinarily be heated above 2300.degree. F. (or 1200.degree. C.), without incurring damage. The usual solution to this overheating problem is to dilute the combustion gases with an excess of compressed air, typically three to four times as much air as is required for stoichiometric combustion with the fuel. This dilution reduces the temperature of the combustion gases below the 2300.degree. F. limit, but unfortunately requires power to compress and deliver the additional air and also significantly increases the engine's size.
An afterburner can be used to fully utilize the dilution air mixed in with the combustion gases. This afterburner is located downstream of the engine's turbine, however, such that the pressure of the air delivered to it has been significantly reduced by passage through the turbine, and the afterburner's efficiency is consequently very low. The afterburner is therefore suitable for use in only very limited circumstances, for relatively short time durations, when additional power is required.
The desire to operate at higher combustion gas temperatures is longstanding. Improved metallurgy and fabrication techniques have permitted operation at somewhat higher temperatures, but there is still significant room for further improvement. Increasing the combustion gas temperature at the turbine inlet merely 600.degree. F., from 1700.degree. F. to 2300.degree. F., nearly doubles the power obtained per pound of air flow.
One technique proposed for permitting engine operation at temperatures of about 2300.degree. F. involves cooling the turbine (shrouding, stator, blades and rotor blades) using air diverted from the engine's compressor. Diverting air from the compressor leads to inefficiencies in the thermal cycle, since a portion of the compressed air (probably about eight percent) bypasses the engine's combustion chamber.
It is well known that the cycle efficiency for the gas turbine with heat recovery goes up when the compression ratio goes down.
Also, it is well known that the pressure ratio for maximum output is higher than that of the pressure ratio for maximum efficiency.
A very important factor which determines the optimum level of the compression ratio is the Cycle Temperature Ratio, ##EQU1##
By raising T', the specific power and the efficiency go up for each optimum pressure ratio.
The maximum T' corresponds with the stoichiometric ratio, which is a theoretically absolute limit for the evolution of the gas turbine.
One well known method to raise the specific work of the turbine cycle is by reheating the gases before they are expanded in the power turbine, which is counter productive in some ways because of the following considerations:
a. More air must be bled from the compressor to cool the hotter power turbine blades, and the effective mass flow through the compressor turbine will be reduced.
b. The practical reheat fuel/air ratio will be far less effective, because the fuel is not burnt in the air but in the combustion gases from the main combustion chamber.
c. The gain in efficiency, due to the reheat obtained with the ideal cycle is not realized in practice, partly because of the additional pressure loss in the reheat chamber.
Today the state-of-the-art technology of the gas turbine is far below the stoichiometric limit.
An example for a modern gas turbine with heat recovery, is defined by the following data: ##EQU2## Pressure ratio=13.75-5.75 Degree of recovery R=0.61-0.66
Thermal efficiency nth=0.2811-0.1915 PA1 Effective efficiency n.sub.e =0.2769-0.1885 PA1 Specific fuel consumption=0.495-0.725 lb/HPh PA1 Specific power=127-48 HP/lb air sec.
It should be appreciated that there is a significant need for a gas turbine engine that avoids the problems identified above, in particular, the poor efficiency and the high specific fuel consumption and the dramatic deterioration of the efficiency at part load.
In particular, there is a need for an engine that combusts fuel and air near the stoichiometric ratio, and that can operate at maximum efficiency for a wide range of loads. The present invention fulfills these and other needs.
The integration of these gas turbines with a power transmission, under the control of a microprocessor programmed to optimize the energetic cycle of the gas turbine with the requirements of the power-user is the final object of this invention.